Compressor shroud with swept grooves

ABSTRACT

A compressor for an aircraft engine. A rotor includes blades rotatable about an axis. Blade tips extend between leading and trailing edges. A shroud surrounds the rotor, with an inner surface surrounding the tips. Grooves are defined in the shroud inner surface adjacent the tips. The grooves extend circumferentially about the shroud and radially from inlet openings to closed end surfaces. Groove sidewalls extend circumferentially about the axis. The grooves are axially spaced-apart, the most upstream inlet opening having an upstream end disposed upstream of the leading edges of the blades. The grooves have a swept angle from the inner surface, with a center of the inlet openings is axially offset of a center of the closed-end surfaces. The grooves span an overall axial distance corresponding to 30% or more of the blades’ chord length. The grooves have circumferential interruptions defined by baffles, and extend non-continuously around a shroud circumference.

TECHNICAL FIELD

The disclosure relates generally to aircraft engines and, moreparticularly, to compressors for such engines.

BACKGROUND

Compressor stall margin is one of many aspects that may affect theoverall performance of aircraft engines. While compressor shrouds orcasings may have various configurations in order to enhance rotor stallmargin, such as surface treatment and/or structural modificationsprovided on the surface of the shroud, minimizing performance loss inthis regard remains desirable.

SUMMARY

There is accordingly provided a compressor for an aircraft engine,comprising: a rotor having a plurality of blades mounted for rotationabout a central axis, the plurality of blades having blade tipsextending between leading and trailing edges; and a shroud surroundingthe rotor and having an inner surface surrounding the blade tips, aplurality of grooves defined in said inner surface of the shroudadjacent said blade tips, the plurality of grooves extendingcircumferentially about the shroud and extending radially from grooveinlet openings defined in the inner surface to closed end surfaces ofthe plurality of grooves, the plurality of grooves having sidewallsextending circumferentially about the central axis, the plurality ofgrooves being axially spaced-apart from each other, the groove inletopening of the most upstream one of the plurality of grooves having anupstream end disposed upstream of the leading edges of the plurality ofblades, the plurality of grooves having a swept angle from the innersurface such that a center of the groove inlet openings is axiallyoffset of a center of a closed-end surface of each of the plurality ofgrooves, the plurality of grooves spanning an overall axial distancecorresponding to 30% or more of a chord length of the plurality ofblades, wherein the plurality of grooves have circumferentialinterruptions defined by a plurality of baffles such that the pluralityof grooves extend non-continuously around a shroud circumference.

There is also provided a compressor for an aircraft engine, comprising:a rotor having a plurality of blades mounted for rotation about acentral axis, the plurality of blades having blade tips extendingbetween leading and trailing edges; and a shroud surrounding the rotorand having an inner surface surrounding the blade tips, a plurality ofgrooves defined in said inner surface of the shroud adjacent said bladetips, the plurality of grooves extending circumferentially about theshroud and extending radially from groove inlet openings defined in theinner surface to closed end surfaces of the plurality of grooves, theplurality of grooves having sidewalls extending circumferentially aboutthe central axis, the plurality of grooves being axially spaced-apartfrom each other, the leading edge of the plurality of blades axiallydisposed between an upstream end of the groove inlet opening of the mostupstream one of the plurality of grooves and a downstream end of thegroove inlet opening of the most upstream one of the plurality ofgrooves, the plurality of grooves having a swept angle from the innersurface such that a center of the groove inlet openings is axiallyoffset of a center of a closed-end surface of each of the plurality ofgrooves, wherein the plurality of grooves have circumferentialinterruptions defined by a plurality of baffles such that the pluralityof grooves extend non-continuously around a shroud circumference.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross sectional view of a gas turbine engine;

FIG. 2 is a schematic cross-sectional view of an exemplary part of thecompressor rotor casing of the engine shown in FIG. 1 ;

FIG. 3 is an enlarged perspective view of an exemplary part of thecompressor rotor casing of FIGS. 1-2 , defining a cross-section A-A anda cross-section B-B;

FIG. 3A is a schematic cross-sectional view taken through A-A in FIG. 3; and

FIG. 3B is a schematic cross-sectional view taken through A-A of analternate compressor rotor casing;

FIG. 4 is another perspective view of the exemplary part of FIG. 3 ,showing the cross-section B-B in a different angle;

FIG. 5 is a schematic cross-sectional view of another exemplary part ofa compressor rotor casing of the engine shown in FIG. 1 ;

FIG. 6 is a side view of an exemplary part of the compressor rotorcasing of FIG. 5 ;

FIGS. 7A-7C are graphical representations of various groove taper anglesin a compressor rotor casing; and

FIGS. 8A-8B are schematic cross-sectional views taken through A-A inFIG. 3 of various groove and baffle configuration options.

DETAILED DESCRIPTION

FIG. 1 illustrates a turbofan gas turbine engine 10 of a type preferablyprovided for use in subsonic flight, generally comprising in serial flowcommunication a transonic fan 12 through which ambient air is propelled,a multistage compressor 14 for pressurizing the air, a combustor 16 inwhich the compressed air is mixed with fuel and ignited for generatingan annular stream of hot combustion gases, and a turbine section 18 forextracting energy from the combustion gases.

The fan 12, also referred to as a low compressor, comprises a rotor 13mounted for rotation about the engine central axis 11. The rotor 13 isprovided with a plurality of radially extending blades 15. Each blade 15has a leading edge 17 and a trailing edge 19 extending radiallyoutwardly from the rotor hub to a tip 21. The rotor 13 is surrounded bya casing 20 including a stationary annular shroud disposed adjacent thetips 21 of the blades 15 and defining an outer boundary for the mainflow path. As shown in FIG. 2 , the casing inner surface is lined with alayer of non-abradable material 22. The layer of non-abradable material22 may thus be considered as part of the casing inner surface, formingpart of the hard shroud wall. In other cases, an abradable material thatmay detach or break from the casing 20 without causing damages, may beused. The radial distance or gap between the tip 21 of the blades 15 andthe adjacent inner surface of the casing 20 is defined as the rotor tipclearance. Each rotor is designed with a nominal rotor tip clearance toprevent or limit interference between the tip 21 of the blades 15 andthe casing 20, which may occur due to rotor imbalance.

Referring to FIG. 2 , it can be seen that a surface treatment is appliedto the low pressure compressor or fan casing 20, though such surfacetreatment may be applied to a high pressure compressor. As will be seenhereinafter, the surface treatment allows stall margin to be increasedand/or tip clearance vortex flow to be weakened and may help to directthe vortex flow in the main flow stream direction. The rotor casingtreatment comprises a series of regularly axially spaced-apartcircumferential grooves 24 defined in the non-abradable region of thecasing inner surface (region of the casing 20 having the layer ofnon-abradable material 22) axially aligned with the tips 21 of theblades 15. Having regularly axially spaced-apart grooves 24, as opposedto irregularly spaced-apart grooves may facilitate manufacturing and/orparametric design of the engine 10 and/or the surface treatment. Inother cases, the grooves 24 may be irregularly or non-uniformly spacedapart in an axial direction along the casing inner surface, as will bediscussed in further detail below.

As shown in FIG. 3 , the grooves 24 do not extend continuously around360 degrees. Stated differently, each groove 24 is intersected orinterrupted over the circumference of the casing 20. In other words, thegrooves 24 have circumferential interruptions such that the grooves 24extend non-continuously around a shroud circumference. In the depictedembodiment, the circumferential interruptions are defined by a pluralityof baffles 30. In other words, each groove 24 comprises a plurality ofsegments 24A extending circumferentially and separated from an adjacentone of the segments 24A by one of the baffles 30. Although not“continuous” along the full circumference of the casing inner surface,each interrupted groove will be referred to as one groove 24 thatcomprises a plurality of groove segments 24A, for simplicity.

In the illustrated example, six shallow circumferentially extendinggrooves 24 are embedded in the non-abradable layer 22 of the rotorshroud around the blades 15. However, it is understood that the seriesof grooves 24 could be composed of more or less than six grooves 24. Forinstance, the rotor casing treatment could comprise from 2 to 15 groovesdepending on the rotor configuration. The grooves 24 may also beirregularly or non-uniformly axially spaced-apart in other embodiments.

Returning to FIG. 2 , in the depicted embodiment, each groove 24 isdefined by a pair of axially opposed sidewalls 26, in this embodimentsubstantially flat, extending forwardly (i.e. towards the front of theengine) from a groove opening (or groove inlet) 25 defined in the shroudsurface 27 to a closed-end surface 28. The closed-end surface 28 may beflat, rounded or semi-circular in various embodiments, as will bediscussed in further detail below. In the depicted embodiment, opposedsidewalls 26 of adjacent grooves 24 intersect at the opening (or“inlet”) 25 with the shroud surface 27, corresponding to a portion ofthe casing inner surface between adjacent grooves 24, forming a sharpedge. Such edge may be rounded up in other embodiments. Illustratively,each opening 25 includes an upstream end 25A and a downstream end 25Brelative to the main flow through the compressor rotor.

As shown in FIG. 2 , each groove 24 has a depth D and a width W. Thegrooves 24 are spaced apart from one another by a spacing X takenaxially along the shroud inner surface 27 (distance between the openingof adjacent grooves 24). Such spacing may be equal between each pair ofaxially adjacent grooves 24. In other cases, the spacing X between afirst pair of axially adjacent grooves 24 may be different, i.e. greateror lesser in magnitude, than the spacing X between another pair ofaxially adjacent grooves 24. Each groove 24 has a depth projection Ynormal to the casing inner surface.

As depicted in FIG. 2 , the groove inlet opening 25 of the first orupstream groove 24 is axially located upstream of the leading edge 17 ofthe blades 15. More particularly, the upstream end 25A of the grooveinlet opening 25 of the first or upstream groove 24 is axially locatedupstream of the leading edge 17 of the blades 15 relative to the mainflow through the compressor rotor. The upstream end 25A is axiallyspaced from the leading edge 17 by a distance L corresponding to, forinstance, 0% to 10% of the chord length of the blades 15. Otherdistances may be contemplated as well. In the shown embodiment, althoughnot necessarily the case in all embodiments, the leading edge 17 of theblades 15 is axially disposed between the upstream end 25A and thedownstream end 25B of the groove inlet opening 25 of the first orupstream groove 24. Other arrangements may be contemplated as well, forinstance both the upstream end 25A and the downstream end 25B of thegroove inlet opening 25 of the first or upstream groove 24 being axiallydisposed upstream of the leading edge 17. In the depicted embodiment,the last or downstream groove 24 is positioned upstream of the bladetrailing edges 19. The grooves 24 may occupy an axial distance ADspanning from the first or upstream groove to the last or downstreamgroove corresponding to 30% or more of the chord length of the blades15. Illustratively, such axial distance AD may be taken from theupstream-most portion of the closed-end surface 28 of the first orupstream groove 24 to the downstream end 25B of the last or downstreamgroove 24. Other reference points for axial distance AD may becontemplated as well. Having the distance L and axial distance AD withinthese ranges may optimize their effect on the flow vortex.

In the shown case, the grooves 24 are forwardly swept (i.e. swepttowards a front of the engine, which may also be upstream relative tothe main gas flow through the compressor rotor) at an angle θ. In otherwords, when viewed axially along the tip 21 of a blade 15 from itsleading edge 17 to its trailing edge 19, such as in FIGS. 2 and 4 , theclosed-end surface 28 of each of the grooves 24 is located upstream ofthe opening 25 of the corresponding groove 24. Alternately defined, thegrooves 24 are inclined such that a center of their inlet openings 25 islocated axially rearward of a center of their closed-end surfaces 28with respect to the orientation of the grooves 24 of the casing 20 inthe engine 10. The angle θ is taken between an axis P normal to thecasing inner surface 27 and a central axis GA extending longitudinallythrough a center of the grooves 24. Angle θ may be referred to as thegroove swept angle, or groove sweep angle, and is more than 0° and lessthan 75°. In an embodiment, the angle θ is at least 10° but no more than75°. Due to the groove swept angle within this range, the swept angledgrooves 24 may contribute to minimizing total pressure loss by havingthe flow exiting from the grooves 24 with a sufficient main flow streamdirection component, and/or may allow maximizing an internal volume ofthe grooves 24 although the layer of non-abradable material 22 of therotor casing may be thin, for maximizing compactness of the rotor casing20 (to reduce weight and/or size of the rotor casing 20). In otherembodiments, the grooves 24 may be rearwardly swept (i.e. swept towardsa rear of the engine, which may also be downstream relative to the maingas flow through the compressor rotor) at an angle θ. In such cases, thegroove swept angle, or groove sweep angle, may be less than 0° and morethan -75° (i.e. a maximum angle of 75° in a rearward direction). In thedepicted embodiment, the grooves are all angled identically, but one ormore of the grooves 24 may have a different angle θ than other ones ormore of the grooves 24 in other embodiments.

In one embodiment, the width W of the grooves 24 is between about 1% toabout 15% of the chord length of the blades 15. The spacing X may haveany suitable value, for instance respecting an aspect ratio X/W is fromabout 0.1 to about 5. Other spacing X between grooves 24 may becontemplated, for instance irregular or uneven distributions. In oneparticular embodiment, the ratio Y/W ranges from about 0.5 to 10. Inmost cases, larger ratios may be better to trap the tip vortex, thoughmanufacturing may limit the possibilities to have a greater ratio (e.g.a ratio greater or much greater than 10).

While in some embodiments the grooves 24 may all have a same geometry,one or more of the grooves may have a respective geometry that maydiffer in one or more dimensions, in some cases.

As shown in FIGS. 2 and 4 , the respective depths D of the grooves 24may vary from the first (most upstream groove 24) to the last, moreparticularly, in this case the respective depths D of the grooves 24increase from the first to the last groove 24, although they may allhave an equal depth D in other embodiments. Depending on theembodiments, the respective depths D of the grooves 24 may substantiallycorrespond to the thickness of the layer of non-abradable material 22 atthe local areas where they are defined. Stated differently, the depthprojection Y of the grooves 24 may substantially correspond to thethickness of the non-abradable material 22. In other cases, the depthsof the grooves 24 may increase or decrease at various rates, or remainconstant, from the first to the last groove 24, as will be discussed infurther detail below.

Now referring to FIG. 3 , the arrays of baffles 30 in the grooves 24 maybe angularly aligned with respect to each other. However, the baffles 30could as well be angularly staggered in the different grooves 24. Inaddition, the number of baffles in the grooves 24 does not have to bethe same. In an embodiment, the number of baffles 30 in each groove 24is greater than the number of rotor blades 15 but less than 5 times ofthe latter. In a particular embodiment, the number of baffles 30 in eachgroove 24 is between 2 and 5 times the number of rotor blades 15. Inanother particular embodiment, there are two times more baffles 30 pergroove 24 than rotor blades 15. Other ratios of baffles 30 per groove 24may be contemplated as well. Having a greater number of baffles 30 pergroove 24 may impede the effects of the casing treatment.

As shown in FIG. 3A, the baffles 30 are provided in the form ofprojections from the closed-end surface 28 of the grooves 24 to theinlet opening 25 thereof. That is, the baffles 30 protrude from theclosed-end surface 28 over a distance corresponding to the full depth Dof the groove 24 in which the baffles 30 are located. The baffles 30 donot necessarily have to be the same shape. The baffles 30 may beintegrally machined, moulded or otherwise formed on the closed-endsurface 28 of the grooves 24. For instance, cutting tools, such asconventional wood ruff cutters, could be used for machining the grooves24 and the baffles 30 in the non-abradable layer 22. In this way, thebaffles 30 can be formed in the grooves 24 in a cost effective manner.The reparability of the casing 20 may be good since the grooves 24 andthe baffles 30 are machined in non-abradable material.

The depicted baffles 30 extend the full width W of the grooves 24between the groove sidewalls 26 (see FIG. 3 ). As shown in FIG. 3 , eachbaffle 30 has a substantially flat surface 32 extending in the sameplane as the shroud inner surface 27. In other words, the flat surface32 of the baffles 30 form a continuous surface with adjacent portions ofthe shroud inner surface. Forming such continuous surface with adjacentportions of the shroud inner surface may contribute to optimizing theeffects of the casing treatment herein described. The flat surface 32may have other shapes, such as concave or other non-flat shape in otherembodiments.

As shown in FIG. 3A, the baffles 30 extends along the full depth D ofthe grooves 24. This may maximize the break of the swirl component(circumferential component) of the main flow stream at the tip of theblades 15 (or simply tip vortex). In the depicted embodiment, thebaffles 30 have two opposed walls 33 spaced apart circumferentially fromeach other and defining respective ends of the baffles 30 (i.e. endsthat are spaced apart in the circumferential direction of the grooves24). In the depicted embodiment, the two opposed walls 33 merge with theflat surface 32 to form a sharp edge at their junction, though roundededges may be contemplated in other embodiments. The grooves closed-endsurface 28 and the baffles 30 form an intersected radially inwardlyfacing surface at the closed end of each groove 24, such that theradially inwardly facing surface is discontinuous along the length(defined along the circumference of the casing inner surface) of eachgroove 24. Although such circumferentially intersected grooves 24 maygenerate flow turbulence due to the baffles 30 opposing thecircumferential component of the tip flow vortex entering and exitingthe grooves 24, such turbulence resulting from the presence of thebaffles 30 may be more beneficial to the performance of the engine 10than if the baffles 30 were omitted entirely, where the circumferentialcomponent of the main flow stream (or tip vortex), would not be suitablycontrolled. The presence of groove interruptions, such as the baffles 30herein described, may enhance the momentum exchanges between main flowand tip clearance flow, hence enhance the effect of the casingtreatment.

Referring to FIG. 3B, another baffle configuration is shown. In thedepicted embodiment, the baffles 30 another embodiment the baffles leanwith an angle ϕ relative to the axis P normal to the casing innersurface 27. In some embodiments, the angle ϕ may vary from -75° to +75°,i.e. into or away from a rotational direction of the blades 15. Theshape of the baffles 30 may vary. For instance, the edges of the bafflesmay be sharp or rounded. A width B of the baffles 30 may be constantalong both radial and axial directions, for instance a tenth of thegroove width W. In other cases, the baffle width B may vary in one orboth of the radial and axial directions. The circumferentialdistribution of baffles may be uniform or uneven, or may assume otherirregular patterns as well.

Referring to FIG. 5 , another exemplary fan casing 20 is shown, withlike reference numerals referring to like elements. The various featuresdiscussed in relation to the fan casing depicted in FIG. 2 may beunderstood to be applicable to the fan casing depicted in FIG. 5 aswell, for instance the upstream end 25A of the groove inlet opening 25of the first or upstream groove 24 being axially located upstream of theleading edge 17 of the blades 15 relative to the main flow through thecompressor rotor. Of note, in the fan casing 20 shown in FIG. 5 , theclosed-end surfaces 28 of the grooves 24 are rounded or semi-circular.Other shapes for the closed-end surfaces 28 may be contemplated as well.In addition, in the embodiment shown in FIG. 5 , the depths D of each isthe grooves 24 is constant from the most upstream groove 24 to the mostdownstream groove 24. Other depths D, for instance increasing ordecreasing depths along the downstream direction, may be contemplated aswell. In the depicted embodiment, the grooves 24 each have a forwardswept angle θ of 45° relative to axis P normal to the casing innersurface 27. Other angles, including rearward swept angles, may becontemplated as well.

Referring to FIG. 6 , the depicted casing 20 includes unevenly-spacedgrooved 24. In other words, spacing X1 between a first pair of grooves24 is different than spacing X2, X3, X4, etc. In the depicted case, theratio between spacing X (X1, X2, X3, X4) and the groove width W (X/W)may vary between 0.5 and 5. In other embodiments, the ratio (X/W) mayvary between 3 and 3.6. Other ratios may be contemplated as well. Asdiscuss above, and in the depicted case, the groove depth D may beconsistent for each groove 24. In the depicted case, each groove 24includes a rounded or semi-circular closed-end surface 28.

Referring to FIGS. 7A-7C, in various embodiments, the taper angle of thegrooves 24, i.e. the variation in radius from one groove 24 to the next,can either remain constant (ex: FIG. 7A), decrease (Ex: FIG. 7B) orincrease (EX: FIG. 7C) from an upstream end to a downstream end of thecasing 20. In FIG. 7A, the taper angle is shown to remain constant, i.e.a taper angle of 0° between grooves 24. In FIG. 7B, an exemplary inwardor decreasing taper angle of 10°, is shown. In FIG. 7C, an exemplaryoutward or increasing taper angle of 10° is shown. Other inward oroutward taper angles may be contemplated. For instance, in various casesthe taper angle may vary from 20° inward to 20° outward.

Referring to FIGS. 8A-8B, the grooves 24 may take on various shapes orpatterns when viewed from cross-section A-A. For instance, the grooves24 depicted in FIG. 7A are shown to have a linearly-circumferentialshape, while the grooves 24 depicted in FIG. 7B are shown to havenon-linear or curved shape. Other groove patterns or shapes, or instancefor instance helically-threaded grooves with baffles, may becontemplated as well.

In the present disclosure, when a specific numerical value is provided(e.g. as a maximum, minimum or range of values), it is to be understoodthat this value or these ranges of values may be varied, for example dueto applicable manufacturing tolerances, material selection, etc. Assuch, any maximum value, minimum value and/or ranges of values providedherein (such as, for example only, the plurality of grooves spanning anoverall axial distance corresponding to 30% or more of a chord length ofthe plurality of blades), include(s) all values falling within theapplicable manufacturing tolerances. Accordingly, in certain instances,these values may be varied by ± 5%. In other implementations, thesevalues may vary by as much as ± 10%. A person of ordinary skill in theart will understand that such variances in the values provided hereinmay be possible without departing from the intended scope of the presentdisclosure, and will appreciate for example that the values may beinfluenced by the particular manufacturing methods and materials used toimplement the claimed technology.

The embodiments described in this document provide non-limiting examplesof possible implementations of the present technology. Upon review ofthe present disclosure, a person of ordinary skill in the art willrecognize that changes may be made to the embodiments described hereinwithout departing from the scope of the present technology. Yet furthermodifications could be implemented by a person of ordinary skill in theart in view of the present disclosure, which modifications would bewithin the scope of the present technology.

1. A compressor for an aircraft engine, comprising: a rotor having aplurality of blades mounted for rotation about a central axis, theplurality of blades having blade tips extending between leading andtrailing edges; and a shroud surrounding the rotor and having an innersurface surrounding the blade tips, a plurality of grooves defined insaid inner surface of the shroud adjacent said blade tips, the pluralityof grooves extending circumferentially about the shroud and extendingradially from groove inlet openings defined in the inner surface toclosed end surfaces of the plurality of grooves, the plurality ofgrooves having sidewalls extending circumferentially about the centralaxis, the plurality of grooves being axially spaced-apart from eachother defining a plurality of axial gaps between adjacent pairs of theplurality of grooves, the groove inlet opening of the most upstream oneof the plurality of grooves having an upstream end disposed upstream ofthe leading edges of the plurality of blades, the plurality of grooveshaving a swept angle from the inner surface such that a center of thegroove inlet openings is axially offset of a center of a closed-endsurface of each of the plurality of grooves, the plurality of groovesspanning an overall axial distance corresponding to 30% or more of achord length of the plurality of blades, wherein the plurality ofgrooves have circumferential interruptions defined by a plurality ofbaffles such that the plurality of grooves extend non-continuouslyaround a shroud circumference.
 2. The compressor as defined in claim 1,wherein the upstream end of the groove inlet opening of the mostupstream one of the plurality of grooves is axially spaced from theleading edge of the plurality of blades by a distance corresponding toat most 10% of the chord length of the plurality of blades.
 3. Thecompressor as defined in claim 1, wherein the plurality of baffles arecircumferentially spaced apart and project from the closed end surfacesto the groove inlet openings.
 4. The compressor as defined in claim 1,wherein the leading edge of the plurality of blades is axially disposedbetween the upstream end of the groove inlet opening of the mostupstream one of the plurality of grooves and a downstream end of thegroove inlet opening of the most upstream one of the plurality ofgrooves.
 5. The compressor as defined in claim 1, wherein a first axialgap of the plurality of axial gaps is defined between a first pair ofadjacent plurality of grooves and a second axial gap of the plurality ofaxial gaps is defined between a second pair of adjacent plurality ofgrooves, the first axial gap having a distance different than a distanceof the second axial gap.
 6. The compressor as defined in claim 5,wherein a ratio of each axial gap distance between pairs of adjacentplurality of grooves and a width of each of the plurality of groovesranges between 0.5 and
 5. 7. The compressor as defined in claim 1,wherein the plurality of grooves have a forwardly swept angle from theinner surface such that the center of the groove inlet openings islocated axially rearward of the center of the closed-end surface of eachof the plurality of grooves.
 8. The compressor as defined in claim 1,wherein each of the plurality of baffles is angled relative to an axisnormal to the inner surface.
 9. The compressor as defined in claim 8,wherein each of the plurality of baffles is angled relative to the axisnormal to the inner surface at an angle ranging from -75 degrees to 75degrees.
 10. The compressor as defined in claim 1, wherein the pluralityof grooves each have a radial depth that increases or decreases inmagnitude from an upstream end of the shroud to a downstream end of theshroud.
 11. The compressor as defined in claim 10, wherein the radialdepth of each of the plurality of grooves increases or decreases at ataper angle of 20 degrees from each of the plurality of grooves to asubsequent downstream groove of plurality of grooves from the upstreamend of the shroud to the downstream end of the shroud.
 12. Thecompressor as defined in claim 1, wherein the closed end surfaces of theplurality of grooves are rounded closed end surfaces.
 13. The compressoras defined in claim 1, wherein the compressor includes a layer ofnon-abradable material lined on the inner surface of the shroud aboutthe blade tips, the layer of non-abradable material embedding theplurality of grooves and baffles.
 14. The compressor as defined in claim1, wherein the grooves have a width between about 1% to about 15% of thechord length of the blades.
 15. The compressor as defined in claim 1,wherein depths of the plurality of grooves are constant from the mostupstream one of the plurality of grooves to the most downstream one ofthe plurality of grooves.
 16. A compressor for an aircraft engine,comprising: a rotor having a plurality of blades mounted for rotationabout a central axis, the plurality of blades having blade tipsextending between leading and trailing edges; and a shroud surroundingthe rotor and having an inner surface surrounding the blade tips, aplurality of grooves defined in said inner surface of the shroudadjacent said blade tips, the plurality of grooves extendingcircumferentially about the shroud and extending radially from grooveinlet openings defined in the inner surface to closed end surfaces ofthe plurality of grooves, the plurality of grooves having sidewallsextending circumferentially about the central axis, the plurality ofgrooves being axially spaced-apart from each other defining a pluralityof axial gaps between adjacent pairs of the plurality of grooves, theleading edge of the plurality of blades axially disposed between anupstream end of the groove inlet opening of the most upstream one of theplurality of grooves and a downstream end of the groove inlet opening ofthe most upstream one of the plurality of grooves, the plurality ofgrooves having a swept angle from the inner surface such that a centerof the groove inlet openings is axially offset of a center of aclosed-end surface of each of the plurality of grooves, wherein theplurality of grooves have circumferential interruptions defined by aplurality of baffles such that the plurality of grooves extendnon-continuously around a shroud circumference.
 17. The compressor asdefined in claim 16, wherein the plurality of grooves span an overallaxial distance corresponding to 30% or more of a chord length of theplurality of blades.
 18. The compressor as defined in claim 16, whereinthe upstream end of the groove inlet opening of the most upstream one ofthe plurality of grooves is axially spaced from the leading edge of theplurality of blades by a distance corresponding to at most 10% of achord length of the plurality of blades.
 19. The compressor as definedin claim 16, wherein a first axial gap of the plurality of axial gaps isdefined between a first pair of adjacent plurality of grooves and asecond axial gap of the plurality of axial gaps is defined between asecond pair of adjacent plurality of grooves, the first axial gap havinga distance different than a distance of the second axial gap.
 20. Thecompressor as defined in claim 16, wherein depths of the plurality ofgrooves are constant from the most upstream one of the plurality ofgrooves to the most downstream one of the plurality of grooves.